The present invention relates generally to air seals for gas turbine engines, and relates more particularly to seals having improved properties in operating conditions during which unusually large amounts of seal material is liberated and ingested into the engine.
Gas turbine engines are well known sources of power, e.g., motive power for aircraft or as power generators, and generally include compressor (typically preceded by one or more fan stages), combustor and turbine sections. As illustrated generally in FIG. 1, compressor and turbine sections (and any fan stages) each include shaft-mounted, rotating disks 1, each carrying a set of blades 2 located within a hollow housing or case 3, with intervening sets of stationary vanes 5 mounted to the case. Air seals 4, 7 are provided between the tips of the blades and the case (outer air seals), and between the vanes and the disks (knife edge seals) to prevent air leakage between those components.
Air is ingested through an engine inlet and compressed by rotating disks and associated blades in the compressor. The compressed air is then burned with fuel in the combustor to generate high pressure and temperature gasses, which cause rotation of the turbine sections and associated fan compressor stages and are then ejected out an engine exhaust to provide thrust. The case is intended to prevent leakage of air or combustion products around the tips of the blades, i.e., between the blade tips and the case, which leakage reduces the efficiency of the engine.
Despite the design of components to minimize leakage, a substantial proportion of any leakage which does occur in a normally-operating gas turbine engine occurs between the tips of the blades and the case, and between the tips of the vanes and the disks. One manner of eliminating such leakage is to fabricate all mating parts to extremely close tolerances, which becomes increasingly expensive as tolerances are reduced. Moreover, given the temperature ranges to which the parts are subjected to before, during and after operation, and the resultant thermal expansion and contraction of the parts, such close tolerances will at times result in interference between mating parts and corresponding component wear and other damage. Accordingly, gas turbine engine designers have devoted significant effort to developing effective air seals, and particularly seals composed of abradable materials. See, e.g., U.S. Pat. No. 4,936,745 to Vine et al. and U.S. Pat. No. 5,706,231 to Nissley et al., which are assigned to the assignee of the present invention and expressly incorporated by reference herein.
Seals require a balance of several properties including relative abradability upon being contacted by a rotating blade tip, erosion resistance, durability, thermal expansion balanced with that of the underlying material, and relative ease and reasonable cost of manufacture. See, e.g., U.S. Pat. No. 5,536,022 to Sileo, which is also assigned to the assignee of the present invention and expressly incorporated by reference herein.
A typical compressor air seal includes the seal substrate, e.g., a metal substrate, an optional metal layer composed of a metal powder plasma sprayed on the substrate, and an abradable, sealing layer applied to the metal layer. Typical sealing layers include a metal matrix of aluminum and silicon with some amount of embedded polyester powder particles and is plasma sprayed onto the substrate, as well as silicone rubber abradable layers incorporating material such as Visilox V-622 from Rhodin of Troy, N.Y. and hollow microspheres. These systems provide adequate performance up to about 500xc2x0 F. While these seal systems have provided adequate performance to date, there remains a desire for a seal system having a higher temperature capability, compatible thermal expansion with the underlying substrate, improved erosion resistance yet readily abrades when contacted by a blade tip of knife edge, and so on.
Moreover, with the desire to reduce the weight of gas turbine engines, particularly for use with aircraft, the use of composite cases for various engine stages has been proposed. In this instance, the use of plasma spray deposition processes is undesirable if not unusable. Accordingly, another type of seal system must be employed.
It is an object of the present invention to provide a gas turbine engine air seal that provides the desired improved performance over present air seals.
It is another object to provide such a seal that is also cost effective.
It is yet another object to provide a seal that weighs no more than conventional seal material, and provides no weight penalty.
It is still another object to provide a seal that can be readily applied to composite substrates.
It is still yet another object to provide such a seal using conventional equipment.
According to one aspect of the present invention, an air seal is disclosed for use in a gas turbine engine. The seal includes a seal substrate, and an abradable seal layer on to the substrate. The abradable layer includes at least a thermoset polymer bulk material such as a phenolic powder and a thermoplastic binder material such as PEEK. The abradable layer may also include a filler to provide some desired characteristic, such as porosity or dry lubrication to enhance abradability.
One advantage of the present invention is that the seal provides improved acceptable durability and abradability, particularly at higher temperatures. In addition, seal of the present invention is cost effective to produce, and does not weigh any more than conventional seal materials.
Additional advantages will become apparent to those skilled in the art ill light of the following description and accompanying drawings.